1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine rotor blade with tip shroud cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, especially an industrial gas turbine engine, a turbine section includes a plurality of rotor blades with stator vanes to direct a hot gas flow through the turbine stages and extract mechanical energy from the hot gas flow. The efficiency of the engine can be increased by passing a higher gas flow into the turbine. However, the factor limiting the highest temperature usable in the turbine is the material properties and the internal cooling ability of the first stage of the turbine. However, the second stage and even the third stage turbine blades and stator vanes can be supplied with cooling air to provide cooling for these airfoils in order to increase the useful life of the parts. Although the cooling requirements of later stage turbine airfoils can usually be easily met, the turbine efficiency can be decreased by using more cooling air than is required. Also, some parts of the turbine airfoils such as the rotor blade tips require cooling at the hot spots. Allowing for excessive hot spots to exist on the airfoils can lead to premature damage or unnecessary creep life damage.
Another method of increasing the efficiency of a turbine is to reduce the leakage that occurs across gaps such as the blade tip gap formed between the rotor blade and the stationary stator casing. Rotor blades make use of an outer shroud member on the radial outer end of the blade. The blade shrouds include abutment faces in which adjacent shrouds form an enclosed flow path for the hot gas flow to pass through the blade stage. The blade shrouds include hard material coatings on the abutting shroud surfaces to increase the useful life of the blades. Leakage across the shroud contact faces will lower the turbine efficiency as well as allow for the high temperature gas flow to affect the hard coatings on the contact faces, leading to creep extension and burning of the coatings and therefore large gaps.
In an industrial gas turbine (IGT) engine (the engine used for electric power generation), the latter stage rotor blades (3rd and 4th stage) are long blades and include shrouds at the blade tips to function as snubbers that dampen vibration found in these larger length blades. The shrouds also form surfaces for the hot gas flow through the turbine stage. With higher temperature turbine inlet temperatures for advanced engines, more cooling capability is required for these blade shrouds. U.S. Pat. No. 5,350,277 issued to Jacal et al on Sep. 27, 1994 and entitled CLOSED-CIRCUIT STEAM-COOLED BUCKET WITH INTEGRALLY COOLED SHROUD FOR GAS TURBINES AND METHODS OF STEAM-COOLING THE BUCKETS AND SHROUDS which discloses a large rotor blade for an IGT with a tip shroud and a knife edge seal that is used to form a hot gas flow seal with an outer shroud of the engine. the tip shroud includes surfaces on both sides that rub against adjacent tip shrouds to dissipate vibrations through friction.
One prior art references attempts to address this problem. U.S. Pat. No. 6,471,480 B1 issued to Balkeum, III et al Oct. 29, 2002 and entitled THIN WALLED COOLED HOLLOW TIP SHROUD discloses a rotor blade tip shroud having cooling air supply passages, metering holes and a plurality of shroud core section to provide cooling for the tip shroud. Cooling holes in the base of the shroud core section also provides cooling air to the tip shroud. Cooling holes are also positioned on the outer walls of the tip shroud core sections to discharge cooling air out from the tip shroud and at the contact surface of the tip shrouds.
U.S. Pat. No. 7,427,188 B2 issued to Neuhoff et al on Sep. 23, 2008 and entitled TURBOMACHINE BLADE WITH FLUIDLY COOLED SHROUD shows another blade tip shroud with cooling. In this tip shroud cooling design, the same cooling air pressure is used throughout the entire tip shroud cooling circuit. Therefore, the various sections of the tip shroud cannot be selectively cooled by passing less or more cooling air to the portions that require less cooling or more cooling.
Another prior art reference provides cooling for the hard contact face of the tip shrouds. U.S. Pat. No. 4,948,338 issued to Wickerson on Aug. 14, 1990 and entitled TURBINE BLADE WITH COOLED SHROUD ABUTMENT SURFACE discloses a tip shroud with a hard face coating being cooled by a wide slot cooling duct. Cooling air through the duct is discharge from the shroud through three ports that are angled downward so that the exhausted cooling air flows over part of the exterior of the coating and also over the part of the exterior of the abutting coating on the adjacent shroud member to provide film cooling for both coatings.
What the two above prior art tip shroud cooling patents do not disclose is the use of impingement cooling for the hard coating on the contact faces of the tip shrouds, or the use of pin fins in the tip shroud cavities or compartments to enhance heat transfer coefficient for improving the cooling of the contact faces and the tip shroud while using less cooling air.